As is well known, a gas turbine engine in its basic form includes a compressor section, a combustion section and a turbine section arranged to provide a generally axially extending flow path for the working gases. Compressed air, from the compressor section, is mixed with fuel and burned in the combustor to add energy to the gases. The hot, pressurized combustion gases are expanded through the turbine section to produce useful work and/or thrust. While an aircraft propulsion engine delivers most of its useful power as forward thrust, other types of gas turbine engines, typically called auxiliary power units, furnish no thrust but are used to supply compressed air and mechanical power to drive electrical generators or hydraulic pumps. The power produced by any engine is a function of, among other parameters, the temperature of the gases admitted into the turbine section. That is, all other things being equal, an increase in power from a given engine can be obtained by increasing the working gas temperature. However, as a practical matter, the maximum gas temperature, and hence the efficiency and output of the engine, is limited by the high temperature capabilities of the various turbine section components in contact with the hot gases.
Within the turbine section are one or more stages of turbine rotor assemblies which are rotated by direct exposure to the hot gases. Such rotors are subjected to very high centrifugal forces and severe thermal gradients as well as high temperatures. There are two basic designs or types of turbine rotors, each having certain operating advantages and disadvantages. An axial-flow wheel has many relatively short, straight, airfoil shaped blades extending radially from the circumference of a generally flat disk mounted on a shaft. Typically the blades are cast individually from one material and mechanically attached to a forged disk of different material so that the properties of each component may be optimized for its particular service environment. The airfoil section of the turbine wheel is susceptible to deformation by high temperature creep and failure by creep rupture induced by the axially directed centrifugal forces imposed upon the blade, and failure by high cycle (low amplitude) fatigue induced by the pulsating impingement of the hot gases. The disks are prone to failure by low cycle fatigue cracking, which can propogate rapidly to burst the disk, caused by very high tensile forces and/or notch sensitivity due to local stress concentrations, either inherent in the disk design or resulting from undetected flaws in critical regions of the disk. Disks and blades now in use have been developed to resist these mechanisms of failure.
The other basic type of turbine rotor, a radial-inflow rotor, presents more challenging design problems when used in a severe operating environment. Radial-flow rotors are generally one piece with a series of thin, scrolled blades or fins arranged in a frusto-conical shape somewhat like a common centrifugal compressor rotor. In operation, hot combustion gases are directed tangentially towards the relatively thin blades near the peripheral rim of the rotor and flow radially inwardly between the blades over a valley-shaped region of the hub surface commonly called the "saddle" before exiting in a generally axial direction.
This hub surface (hereinafter saddle) is an area of high mechanical stress concentration due to the geometry of the rotor. In addition, it is rapidly heated by the hot gas whereas the interior of the hub responds more slowly during a cold engine start. Thus, a transient thermal gradient is created within the saddle region which causes extreme circumferential compression at or near the saddle surface. When the engine is unloaded or shutdown, the hot gas temperature rapidly drops to a lower level. This reverses the thermal gradient by rapidly cooling the saddle, thus producing circumferential tension which adds to the tensile stresses produced by centrifugal forces. Such subjection of the saddle to high temperature compression and subsequent rapid cooling contraction creates structural cracks, thought to be due to low-cycle thermal/mechanical fatigue, which propogates into the hub and can lead to eventual destruction of the entire turbine rotor.
Several different approaches to solving or reducing this cracking problem have been tried by prior researchers in this field in order to extend the useful life of such rotors.
One early attempt involved the addition of cooling air passageways within the turbine rotor adjacent the hot front face in order to reduce the maximum temperatures in the saddle. See, for example, U.S. Pat. Nos. 2,873,945 and 4,587,700. However, such internal passageways are difficult or costly to manufacture, require a complex ducting system to supply the cooling air to the rotor and generally result in high local concentrations of stress which may lead to failure.
Another approach has been to refine the geometric design of the rotor to minimize as much as possible, the peak stresses due to the variations in temperature. Even though the stress and temperature conditions within radial turbines, especially in the saddle area, are very difficult to characterize due to the complex geometry and the thermal boundary conditions, some improvements have been developed by trial-and-error testing. For example, it was discovered that the life of radial turbines could be improved significantly, with only a small loss in aerodynamic efficiency, by removing the thinnest material between the blades so that the saddle regions were more massive and thus more resistant to thermal cracking. Such scalloped rotors are now common in the art but further improvements are necessary to meet the increasing demands for longer life or higher performance from turbine engines.
One of the more recent approaches has been to construct rotors of two (or more) different materials so that each portion of the rotor will have its metallurgical properties optimized for its local operating conditions as was done for axial-flow wheels. That is, the hub is typically forged from vacuum melted ingots (or are consolidated from fine alloy powders) to have superior low temperature tensile strength and low cycle fatigue (LCF) resistance but quite limited high temperature creep rupture properties. On the other hand, the blade portion is typically cast from a relatively fine grain alloy which has the reverse properties, i.e. good creep strength but relatively poor tensile and LCF proprties. It has even been proposed to form portions of the blade tips from directionally solidified or single crystal superalloys developed for axial-flow turbine wheels but such approaches have not been commercially successful. The problems of adequately bonding the two portions have, for the most part, been solved in the prior art by hot isostatic pressing and/or diffusion bonding. For a more complete discussion of the manufacture and benefits of dual alloy rotors see, for example, U.S. Pat. Nos. 4,335,997, 4,581,300: 4,529,452, 4,659,288, and 4,787,821 which are incorporated herein by reference.
Even though the theoretical operating life of the blade and hub portions have been increased by these prior approaches (either alone or in combination) the overall useful life of a rotor is still limited by cracking in the saddle regions to a value about an order of magnitude lower than either of the components. Further improvements in the cyclic fatigue resistance of the components are not resulting in a significant improvement in a rotors useful life. It is believed that the thermal stress range in the saddle region needs to be reduced, preferably without sacrificing aerodynamic performance.
In view of the foregoing needs of the art, it is an object of the present invention to provide a novel radial flow turbine rotor having improved LCF life in the saddle regions by reducing the effects of cyclic thermal stresses therein.
Another object of the invention is to provide a method of manufacturing a radial flow rotor having an increased resistance to saddle cracking without a sacrifice in aerodynamic performance.
Further objects and features of the invention should become apparent from the following specification and claims.